US8371816B2 - Rotor blades for turbine engines - Google Patents

Rotor blades for turbine engines Download PDF

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Publication number
US8371816B2
US8371816B2 US12/533,378 US53337809A US8371816B2 US 8371816 B2 US8371816 B2 US 8371816B2 US 53337809 A US53337809 A US 53337809A US 8371816 B2 US8371816 B2 US 8371816B2
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Prior art keywords
damping fin
tip shroud
leading edge
trailing edge
edge damping
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Expired - Fee Related, expires
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US12/533,378
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English (en)
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US20110027088A1 (en
Inventor
Matthew R. Piersall
Brian D. Potter
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GE Vernova Infrastructure Technology LLC
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General Electric Co
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Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: PIERSALL, MATTHEW R., POTTER, BRIAN D.
Priority to US12/533,378 priority Critical patent/US8371816B2/en
Priority to JP2010117907A priority patent/JP5681384B2/ja
Priority to DE102010017105A priority patent/DE102010017105A1/de
Priority to CH00852/10A priority patent/CH701537B1/de
Priority to CN2010102004441A priority patent/CN101988392A/zh
Publication of US20110027088A1 publication Critical patent/US20110027088A1/en
Assigned to U.S. DEPARTMENT OF ENERGY reassignment U.S. DEPARTMENT OF ENERGY CONFIRMATORY LICENSE (SEE DOCUMENT FOR DETAILS). Assignors: GENERAL ELECTRIC COMPANY
Publication of US8371816B2 publication Critical patent/US8371816B2/en
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Assigned to GE INFRASTRUCTURE TECHNOLOGY LLC reassignment GE INFRASTRUCTURE TECHNOLOGY LLC ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: GENERAL ELECTRIC COMPANY
Expired - Fee Related legal-status Critical Current
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/16—Form or construction for counteracting blade vibration
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/96—Preventing, counteracting or reducing vibration or noise

Definitions

  • the present application relates generally to apparatus, methods and/or systems concerning the design and operation of turbine rotor blades. More specifically, but not by way of limitation, the present application relates to apparatus, methods and/or systems pertaining to turbine blade tip shrouds with damping and other features.
  • a gas turbine engine In a gas turbine engine, it is well known that air pressurized in a compressor is used to combust a fuel in a combustor to generate a flow of hot combustion gases, whereupon such gases flow downstream through one or more turbines so that energy can be extracted therefrom.
  • rows of circumferentially spaced turbine rotor blades extend radially outwardly from a supporting rotor disk.
  • Each blade typically includes a dovetail that permits assembly and disassembly of the blade in a corresponding dovetail slot in the rotor disk, as well as an airfoil that extends radially outwardly from the dovetail and interacts with the flow of the working fluid through the engine.
  • the airfoil has a generally concave pressure side and generally convex suction side extending axially between corresponding leading and trailing edges and radially between a root and a tip. It will be understood that the blade tip is spaced closely to a radially outer turbine shroud for minimizing leakage therebetween of the combustion gases flowing downstream between the turbine blades.
  • rotor blades due to various stimulus sources during engine operation, rotor blades often exist in a state of vibration or resonance.
  • the sources of vibration generally include rotational imbalance, stator blade stimulus, unsteady pressure perturbations, and combustion acoustic tones.
  • the resulting vibration generally results in the accrual of high cycle fatigue damage, which typically shortens the life of the rotor blade and, in cases where the fatigue causes a blade failure during operation, may lead to catastrophic damage to the turbine engine.
  • the magnitude of the vibration is related at least in part to the amount of damping that is introduced into the system. The more damping that is introduced, the lower the vibratory response, and the more reliable the turbine system becomes. As such, there is a continuing need for improved apparatus, system, and methods for damping and, thereby, reducing the vibration experienced by the rotor blades of turbine engine during operation.
  • the present application thus describes a tip shroud that includes a plurality of damping fins, each damping fin comprising a substantially non-radially-aligned surface that is configured to make contact with a tip shroud of a neighboring rotor blade.
  • At least one damping fin comprises a leading edge damping fin and at least one damping fin comprise a trailing edge damping fin; and the leading edge damping fin corresponds to the trailing edge damping fin.
  • the present application further describes a tip shroud for a turbine rotor blade that includes a plurality of damping fins, each damping fin comprising a substantially non-radially-aligned contact surface that is configured to make contact with a tip shroud of a neighboring rotor blade.
  • At least one damping fin may comprise a leading edge damping fin and at least one damping fin may comprise a trailing edge damping fin.
  • the leading edge damping fin and the trailing edge damping fin may be configured such that when a set of rotor blades having tip shrouds of the same design are installed in a rotor disk of the turbine engine, the leading edge damping fin of a first rotor blade engages the trailing edge damping fin of a second rotor blade that directly leads the first rotor blade and the trailing edge damping fin of the first rotor blade engages the leading edge damping fin of a third rotor blade that directly trails the first rotor blade.
  • the radial position of the leading edge damping fin may be offset from the radial position of the trailing edge damping fin such that a desired level of contact between the substantially non-radially-aligned contact surface of the leading edge damping fin and the substantially non-radially-aligned contact surface of the trailing edge damping fin is maintained during operation of the turbine engine.
  • FIG. 1 is a schematic representation of an exemplary gas turbine engine in which embodiments of the present application may be used;
  • FIG. 2 is a sectional view of the compressor in the gas turbine engine of FIG. 1 ;
  • FIG. 3 is a sectional view of the turbine in the gas turbine engine of FIG. 1 ;
  • FIG. 4 is a perspective view of an exemplary gas turbine engine rotor blade having a tip shroud of conventional design
  • FIG. 5 is an outboard view of a series of installed turbine blades having a tip shrouds of conventional design
  • FIG. 6 is a perspective view of the leading edge of a turbine engine rotor blade having a tip shroud and a damping fin according to an exemplary embodiment of the present application;
  • FIG. 7 is a perspective view of the trailing edge of the turbine engine rotor of FIG. 6 having a tip shroud and corresponding damping fin according to an exemplary embodiment of the present application.
  • FIG. 8 is a perspective view of the leading edge of a turbine engine rotor blade having a tip shroud according to an exemplary embodiment of the present application and, more particularly, possible angular configurations for a damping fin according to the present application.
  • rotor blade without further specificity, is a reference to the rotating blades of either the compressor 52 or the turbine 54 , which include both compressor rotor blades 60 and turbine rotor blades 66 .
  • stator blade without further specificity, is a reference the stationary blades of either the compressor 52 or the turbine 54 , which include both compressor stator blades 62 and turbine stator blades 68 .
  • blades will be used herein to refer to either type of blade.
  • blades is inclusive to all type of turbine engine blades, including compressor rotor blades 60 , compressor stator blades 62 , turbine rotor blades 66 , and turbine stator blades 68 .
  • downstream and upstream are terms that indicate a direction relative to the flow of working fluid through the turbine.
  • downstream refers to a direction that generally corresponds to the direction of the flow of working fluid
  • upstream generally refers to the direction that is opposite of the direction of flow of working fluid.
  • the terms “trailing” and “leading” generally refers relative position in relation to the direction of rotation for rotating parts.
  • the “leading edge” of a rotating part is the front or forward edge given the direction that the part is rotating and, the “trailing edge” of a rotating part is the aft or rearward edge given the direction that the part is rotating.
  • the term “radial” refers to movement or position perpendicular to an axis. It is often required to described parts that are at differing radial positions with regard to an axis. In this case, if a first component resides closer to the axis than a second component, it may be stated herein that the first component is “radially inward” or “inboard” of the second component.
  • first component resides further from the axis than the second component, it may be stated herein that the first component is “radially outward” or “outboard” of the second component.
  • axial refers to movement or position parallel to an axis.
  • circumferential refers to movement or position around an axis.
  • FIGS. 1 through 3 illustrate an exemplary gas turbine engine in which embodiments of the present application may be used. It will be understood by those skill in the art that the present invention is not limited to this type of usage. As stated, the present invention may be used in gas turbine engines, such as the engines used in power generation and airplanes, steam turbine engines, and other type of rotary engines.
  • FIG. 1 is a schematic representation of a gas turbine engine 50 .
  • gas turbine engines operate by extracting energy from a pressurized flow of hot gas produced by the combustion of a fuel in a stream of compressed air.
  • gas turbine engine 50 may be configured with an axial compressor 52 that is mechanically coupled by a common shaft or rotor to a downstream turbine section or turbine 54 , and a combustor 56 positioned between the compressor 52 and the turbine 56 .
  • FIG. 2 illustrates a view of an exemplary multi-staged axial compressor 52 that may be used in the gas turbine engine of FIG. 1 .
  • the compressor 52 may include a plurality of stages. Each stage may include a row of compressor rotor blades 60 followed by a row of compressor stator blades 62 .
  • a first stage may include a row of compressor rotor blades 60 , which rotate about a central shaft, followed by a row of compressor stator blades 62 , which remain stationary during operation.
  • the compressor stator blades 62 generally are circumferentially spaced one from the other and fixed about the axis of rotation.
  • the compressor rotor blades 60 are circumferentially spaced and attached to the shaft; when the shaft rotates during operation, the compressor rotor blades 60 rotate about it. As one of ordinary skill in the art will appreciate, the compressor rotor blades 60 are configured such that, when spun about the shaft, they impart kinetic energy to the air or fluid flowing through the compressor 52 .
  • the compressor 52 may have other stages beyond the stages that are illustrated in FIG. 2 . Additional stages may include a plurality of circumferential spaced compressor rotor blades 60 followed by a plurality of circumferentially spaced compressor stator blades 62 .
  • FIG. 3 illustrates a partial view of an exemplary turbine section or turbine 54 that may be used in the gas turbine engine of FIG. 1 .
  • the turbine 54 also may include a plurality of stages. Three exemplary stages are illustrated, but more or less stages may present in the turbine 54 .
  • a first stage includes a plurality of turbine buckets or turbine rotor blades 66 , which rotate about the shaft during operation, and a plurality of nozzles or turbine stator blades 68 , which remain stationary during operation.
  • the turbine stator blades 68 generally are circumferentially spaced one from the other and fixed about the axis of rotation.
  • the turbine rotor blades 66 may be mounted on a turbine wheel (not shown) for rotation about the shaft (not shown).
  • a second stage of the turbine 54 also is illustrated.
  • the second stage similarly includes a plurality of circumferentially spaced turbine stator blades 68 followed by a plurality of circumferentially spaced turbine rotor blades 66 , which are also mounted on a turbine wheel for rotation.
  • a third stage also is illustrated, and similarly includes a plurality of turbine stator blades 68 and rotor blades 66 .
  • the turbine stator blades 68 and turbine rotor blades 66 lie in the hot gas path of the turbine 54 .
  • the direction of flow of the hot gases through the hot gas path is indicated by the arrow.
  • the turbine 54 may have other stages beyond the stages that are illustrated in FIG. 3 .
  • Each additional stage may include a row of turbine stator blades 68 followed by a row of turbine rotor blades 66 .
  • the rotation of compressor rotor blades 60 within the axial compressor 52 may compress a flow of air.
  • energy may be released when the compressed air is mixed with a fuel and ignited.
  • the resulting flow of hot gases from the combustor 56 which may be referred to as the working fluid, is then directed over the turbine rotor blades 66 , the flow of working fluid inducing the rotation of the turbine rotor blades 66 about the shaft.
  • the mechanical energy of the shaft may then be used to drive the rotation of the compressor rotor blades 60 , such that the necessary supply of compressed air is produced, and also, for example, a generator to produce electricity.
  • FIGS. 4 and 5 illustrate a tip shrouded turbine rotor blade 100 according to conventional design.
  • the turbine rotor blade 100 includes a dovetail 101 which may have any conventional form, such as an axial dovetail configured for being mounted in a corresponding dovetail slot in the perimeter of the rotor disk.
  • An airfoil 102 is integrally joined to the dovetail 101 and extends radially or longitudinally outwardly therefrom.
  • the rotor blade 100 also includes a platform 103 disposed at the junction of the airfoil 102 and the dovetail 101 for defining a portion of the radially inner flowpath through the turbine engine.
  • the airfoil 102 is the active component of the blade 100 that intercepts the flow of the working fluid.
  • a tip shroud 104 may be positioned at the top of the airfoil 102 .
  • the tip shroud 104 essentially is an axially and circumferentially extending flat plate that is supported towards its center by the airfoil 102 .
  • Positioned along the top of the tip shroud 104 may be a seal rail 106 .
  • the seal rail 106 projects radially outward from the outer radial surface of the tip shroud 104 .
  • the seal rail 106 generally extends circumferentially between opposite ends of the tip shroud in the general direction of rotation.
  • the seal rail 106 is formed to deter the flow of working fluid through the gap between the tip shroud 104 and the inner surface of the surrounding stationary components.
  • the seal rails 106 extend into an abradable stationary honeycomb shroud that opposes the rotating tip shroud 104 .
  • a cutter tooth 107 may be disposed toward the middle of the seal rail 106 so as to cut a groove in the honeycomb of the stationary shroud that is slightly wider than the width of the seal rail 106 .
  • Tip shrouds 104 may be formed such that the tip shrouds 104 of neighboring blades make contact during operation.
  • FIG. 5 illustrates an outboard view of turbine rotor blades as they might appear when assembled on a turbine rotor disk and provides an example of a conventional arrangement where neighboring tip shrouds 104 make contact with each other during operation. Two full neighboring tip shrouds are shown with an arrow indicating the direction of rotation. As depicted, the trailing edge of the leading tip shroud 104 may contact or come in close proximity to the leading edge of the trailing tip shroud 104 .
  • This area of contact is often generally referred to as an interface or contact face 108 , or, more particularly, given the configuration of the example provided, a Z-interface 108 .
  • the Z-interface 108 may be so-named because of the approximate “Z” shaped profile between the two edges of the neighboring tip shrouds 104 .
  • the use of the turbine blade 100 and the tip shroud 104 are exemplary only and that other turbine blades and tip shrouds of different configurations may be used with alternative embodiments of the current application. Further, the use of a “Z” shaped interface is exemplary only.
  • a narrow space may exist at the contact face (or Z-interface) 108 between the edges of adjacent tip shrouds 104 .
  • the expansion of the turbine blade metal and the “untwist” of the airfoil may cause the gap to narrow such that the edges of adjacent tip shrouds 104 make contact.
  • Other operating conditions including the high rotation speeds of the turbine and the related vibration, may cause contact between adjacent tip shrouds 104 , even where a gap in the contact face 108 partially remains during turbine operation.
  • One of the functions of the contact made between neighboring tip shrouds 104 is to damp the system and, thereby, reduce vibration.
  • FIGS. 6 and 7 illustrate an exemplary embodiment of the claimed invention, a tip shroud 200 .
  • FIG. 6 illustrates the leading edge of the tip shroud 200
  • FIG. 7 illustrates the trailing edge.
  • the tip shroud 200 may have a first contact surface or radially-aligned contact surface 202 .
  • the radially-aligned contact surface 202 refers to one or more contact surfaces (i.e., surfaces configured to make contact with the tip shrouds of neighboring rotor blades) that are aligned approximately in the radial direction. As one of ordinary skill in the art will appreciate, this primarily includes the surface toward the middle of the tip shroud 200 that extends radially outward along the seal rail 106 .
  • the radially-aligned contact surface 202 may also include any radially-aligned contact surfaces, including those that extend outward from the middle of the tip shroud 200 along the axial length of the tip shroud 200 .
  • the radially-aligned contact surfaces include surfaces that are substantially aligned in the radial direction.
  • a radially-aligned contact surface may be a surface within +/ ⁇ 10 degrees of a radial reference line.
  • the tip shroud 200 also may include a substantially non-radially-aligned second contact surface that is formed via a protrusion from the tip shroud 200 , which herein is referred to as a “damping fin 204 .”
  • the damping fin 204 may include a fin or tab type protrusion that extends substantially both circumferentially and axially from either the leading or trailing edge of the tip shroud 200 .
  • the damping fin 204 may have a relatively narrow or thin profile.
  • the damping fin 204 may extend or slope in a radial direction as well. In those type of embodiments, as defined in more detail below, the extent of the damping fin 204 radial slope will be substantially less steep than the radially aligned contact surface 202 described above.
  • one of the damping fins 204 may be located on the leading edge of the tip shroud 200 , and, as shown in FIG. 7 , another damping fin 204 may be positioned on the trailing edge of the tip shroud 200 .
  • the leading edge damping fin 204 may be located on the pressure side of the tip shroud 200
  • the trailing edge damping fin 204 may be located on the suction side of the tip shroud 200 , though, other configurations, as explained in more detail below, are also possible.
  • the damping fins 204 on the leading and trailing edges of the tip shroud 200 may be configured to correspond with each other.
  • damping fins “corresponding” is intended to mean that when a set of rotor blades having tip shrouds of the same design are properly installed in a rotor disk of a turbine engine, the damping fin 204 positioned on the leading edge of the tip shroud 200 of a first rotor blade (i.e., a “leading edge damping fin”) resides in a desired position in relation to the damping fin 204 positioned on the trailing edge of the tip shroud 200 of a second rotor blade (i.e., a “trailing edge damping fin”) that trails the first rotor blade.
  • damping fins “corresponding” also means that the trailing edge damping fin 204 of the first rotor blade resides in a desired position in relation to the leading edge damping fin 204 of a third rotor blade that leads the first rotor blade.
  • the corresponding damping fins 204 may engage each other. In other embodiments, the corresponding damping fins 204 may reside in close proximity to each other.
  • the radial position of the leading edge damping fin 204 and the trailing edge damping fin 204 may be offset slightly so to produce the desired level of contact or proximity between the corresponding trailing edge damping fin and the leading edge damping fin during operation.
  • the corresponding damping fins 204 may reside in close radial position to each other and, having similar size and shape, may be configured such that the corresponding damping fins 204 of neighboring rotor blades substantially overlap each other axially and circumferentially.
  • the extent of the radial offset may determine the amount of contact made during operation.
  • the radial offset is configured such that the contact surfaces of corresponding damping fins 204 touch or engage each other.
  • the radial offset is configured such that the contact surfaces of corresponding damping fins 204 do not touch each other when they turbine is “cold” or during engine startup (i.e., a startup phase), but make regular contact as the engine warms during operation thereafter.
  • the radial offset is configured such that the contact sources of corresponding damping fins 204 do not touch each other when the turbine engine is “cold” or during engine startup, but make partial contact as the engine warms during operation.
  • the radial offset is configured such that the contact surfaces of corresponding damping fins 204 make partial contact when the turbine engine is “cold” or during engine startup, but make relatively constant contact as the engine warms during operation.
  • the trailing edge damping fin 204 may be positioned just outboard of the leading edge damping fin 204 .
  • a contact face is formed on the outer radial surface of the leading edge damping fin 204 .
  • a contact face is formed on the inner radial surface of the trailing edge damping fin 204 .
  • such contact faces may be provided with enhanced wear properties to prolong the life of the part.
  • the contact face may be provided with a wear coating or more durable material.
  • the contact faces are formed with a cobalt-based hardfacing powder.
  • the damping fins 204 may be configured such that during turbine engine operation, the outer radial surface of the leading edge damping fin 204 and the inner radial surface of the trailing edge damping fin 204 of adjacent turbine blades make at least partial contact. This contact, as one of ordinary skill in the art will appreciate, generally mechanically dampens some of the vibration being experienced by the rotor blades.
  • the damping fin 204 may have an approximate rectangular shape that includes somewhat rounded corners, as shown. Other shapes are possible, including semicircular. Further, while a preferred embodiment is shown in FIGS. 6 and 7 , other arrangements and configurations are possible. For example, in another preferred embodiment, the leading edge damping fin may be positioned on the suction side of the tip shroud and the trailing edge damping shroud may be positioned on the pressure side of the tip shroud. In addition, the leading edge damping fin, instead of being positioned inboard, may be position outboard of the trailing edge damping fin.
  • the trailing edge damping fin may include fins on both the pressure side and suction side of the tip shroud, and the leading edge damping fins may include damping fins that correspond to these on both the pressure side and suction side of the tip shroud.
  • the leading edge damping fins may be inboard, outboard, or both inboard and outboard in relation to the corresponding trailing edge damping fins. More particularly, in one embodiment, one of the leading edge damping fins may be inboard of a corresponding trailing edge damping fin, while the other leading edge damping fin is outboard of the corresponding trailing edge damping fin. In some applications, this interlocking configuration may provide enhanced damping characteristics.
  • the damping fins 204 are configured such that the fins extend primarily circumferentially and axially. That is, the damping fins 204 form an angle with the radial direction of the turbine engine of approximately 90 degrees, and, accordingly, as shown, the damping fins 204 form an angle with the axial direction and the circumferential direction of the turbine engine of approximately 0 degrees. In some embodiments, this angle or slope may be adjusted or tuned to increase the damping of a single vibration mode or several different vibration modes that might be particularly troublesome or heretofore unaffected by other conventional damping efforts, as one of ordinary skill in the art will appreciate. In this manner, the secondary contact surface, i.e., the damping fin 204 , may be designed to provide damping for a vibration mode that might not have been adequately addressed by a conventional radially-aligned damping contact surface.
  • FIG. 8 illustrates how the angle of the damping fin 204 may be adjusted such that different vibration modes may be addressed. As shown, in one embodiment, this may be accomplished by rotating the damping fin 204 about an axis formed at the base of the damping fin, i.e., where the damping fin 204 protrusion connects to the tip shroud 200 . In this manner, the modes of vibration that are dampened by the damping fin 204 may be manipulated in a desired manner. If one of the damping fins 204 is rotated, it will be appreciated that the corresponding damping fin 204 at the other edge of the tip shroud will be oppositely rotated to substantially the same angle.
  • the damping fins 204 may still make contact along a significant or substantially all of their respective contact surfaces.
  • the configuration of damping fins includes at least one trailing edge damping fin on both the pressure side and suction side of the tip shroud and at least one leading edge damping fin on both the one pressure side and suction side of the tip shroud.
  • each of leading edge damping fins corresponds to one of the trailing edge damping fins.
  • the angle of rotation of the damping fin 204 may vary depending on the application.
  • the angle of rotation of the damping fin 204 may be identified generally by the angle the damping fin 204 makes with a radially oriented reference line.
  • the damping fins 204 forms an angle with the radial reference line of approximately 90 degrees.
  • the damping fins may form an angle with the radial reference line of between approximately 70 and 110 degrees.
  • the damping fins may form an angle with the radial reference line of between approximately 60 and 120 degrees.
  • the damping fins may form an angle with the radial reference line of between approximately 45 and 135 degrees.
  • the damping fins may form an angle with the radial reference line of between approximately 30 and 150 degrees.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US12/533,378 2009-07-31 2009-07-31 Rotor blades for turbine engines Expired - Fee Related US8371816B2 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US12/533,378 US8371816B2 (en) 2009-07-31 2009-07-31 Rotor blades for turbine engines
JP2010117907A JP5681384B2 (ja) 2009-07-31 2010-05-24 タービンエンジン用のロータブレード
DE102010017105A DE102010017105A1 (de) 2009-07-31 2010-05-27 Rotorschaufeln für Turbinenanlagen
CH00852/10A CH701537B1 (de) 2009-07-31 2010-05-28 Spitzendeckplatte mit Dämpfungsrippen für eine Rotorschaufel, die in eine Rotorscheibe einer Turbinenanlage einsetzbar ist.
CN2010102004441A CN101988392A (zh) 2009-07-31 2010-05-31 用于涡轮发动机的转子叶片

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US12/533,378 US8371816B2 (en) 2009-07-31 2009-07-31 Rotor blades for turbine engines

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US20110027088A1 US20110027088A1 (en) 2011-02-03
US8371816B2 true US8371816B2 (en) 2013-02-12

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US (1) US8371816B2 (enrdf_load_stackoverflow)
JP (1) JP5681384B2 (enrdf_load_stackoverflow)
CN (1) CN101988392A (enrdf_load_stackoverflow)
CH (1) CH701537B1 (enrdf_load_stackoverflow)
DE (1) DE102010017105A1 (enrdf_load_stackoverflow)

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US20170356298A1 (en) * 2016-06-08 2017-12-14 Rolls-Royce Plc Stator vane
US20180010467A1 (en) * 2016-07-06 2018-01-11 General Electric Company Shroud configurations for turbine rotor blades

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CN102877892B (zh) * 2012-10-23 2015-02-11 湖南航翔燃气轮机有限公司 一种涡轮转子叶片及具有其的燃气轮机
PL416301A1 (pl) * 2016-02-29 2017-09-11 General Electric Company Zespół bandaża silnika turbinowego
CN110312846B (zh) 2017-02-23 2022-05-10 三菱动力株式会社 涡轮动叶以及燃气轮机
FR3075282B1 (fr) * 2017-12-14 2021-01-08 Safran Aircraft Engines Dispositif amortisseur
CN111615584B (zh) * 2017-12-18 2022-08-16 赛峰飞机发动机公司 阻尼装置
CN109057871A (zh) * 2018-04-20 2018-12-21 西门子（中国）有限公司 汽轮机叶冠及叶冠单元
CN112313395B (zh) 2018-06-19 2023-03-07 三菱重工业株式会社 涡轮动叶、涡轮机械以及接触面制造方法
US11053804B2 (en) * 2019-05-08 2021-07-06 Pratt & Whitney Canada Corp. Shroud interlock
US11085303B1 (en) * 2020-06-16 2021-08-10 General Electric Company Pressurized damping fluid injection for damping turbine blade vibration
US11608747B2 (en) * 2021-01-07 2023-03-21 General Electric Company Split shroud for vibration reduction

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JP5681384B2 (ja) 2015-03-04
US20110027088A1 (en) 2011-02-03

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